A modern gas turbine engine, such as is used for generation of electricity at power plants, is a multi-part assembly of various adjacent components, many of which are subjected to mechanical and thermal stresses over long periods of operation. Mechanical stresses to various components result from one or more of vibrational, low cycle (and other thermally related), and other types of stress contributors. As to direct thermal stresses, operating temperatures in some gas turbine engine combustion chambers may reach or exceed 2,900 degrees Fahrenheit, and components of such combustion chambers are cooled and/or provided with thermal barrier coatings to address exposure to such elevated temperatures.
When cooling is used to maintain a component below a specified temperature, often compressed air from the compressor is diverted to pass through a cooling passage. In “closed cooling” approaches, such air, after cooling, continues to the combustor entrance and joins the major flow of air supplied for combustion and dilution purposes. In “open cooling” approaches, such air enters a hot gas flow path downstream of the combustor and may be less available, or not available, for such primary purposes.
Also, as operating temperatures are elevated, there is a greater concern for effective dilution of the fuel/air combustion mixture to lower the NOx to reach desired emissions standards. As compressed air is used for cooling components, such as through open cooling of more downstream components, this loss of compressed air that would otherwise enter the combustor may result in higher than desired NOx emissions.
The junction between the combustor (which generally may be considered to comprise a combustion chamber) and the transition of a gas turbine engine typically has a spring clip assembly that provides for a relatively tight but flexible connection between these components. This connection provides for the combustor/transition assembly to expand and contract as needed, relative to the outer casing, during thermal changes, while also providing a seal between the combustor and the transition. The prior art spring clip assemblies are designed to allow air to flow through such spring clip assemblies and this provides an open type cooling to the spring clips and adjacent components. However, the level and variability of cooling air flow through various existing spring clip assemblies does not provide a level of precision and accuracy for cooling, and consequent cooling efficiency, that is desired for more advanced gas turbine engine systems.
There have been various efforts to improve aspects of the seal between a combustor and a transition of a gas turbine engine. For example, U.S. Pat. No. 6,869,082, issued Mar. 22, 2005 to David M. Parker, teaches an improved spring clip seal in which at least one leaf may include a flared end for limiting gas from passing through slots in the seal of the spring clip. A center sealing member is positioned in at least one embodiment between inner and outer spring clip housings. U.S. Pat. No. 7,007,482, issued Mar. 7, 2006 to A. Green et al., teaches an alternate interface region between a combustion liner and a transition duct. This region comprises feed holes supplying cooling fluid into an annulus and a means for augmenting heat transfer which may comprise geometric ridge configurations. These are stated to help achieve a heat transfer augmentation by turbulating the cooling air to maximize the cooling effectiveness.
Notwithstanding such efforts, a need remains for a seal between a combustor and a transition that provides for more precise and accurate cooling.